Although not limited thereto, the invention is described in its applications to a helicopter of the hybrid type, e.g. as described in documents FR 2 916 418, FR 2 916 419, and FR 2 916 420.
That hybrid helicopter associates, at reasonable cost, the effectiveness in vertical flight of a conventional helicopter with high-speed performance in horizontal flight, e.g. up to a cruising air speed of about 400 kilometers per hour (km/h) or even more, and indeed also with a greater range.
For reasons of simplicity, the term “rotary wing” is used below to designate the lift rotor(s) of such an aircraft, and the term “rotor disk” is used to designate the geometrical surface defined by such a rotor.
Unlike a conventional helicopter, a hybrid helicopter has not only its rotary wing, but also a fixed wing that provides part of its lift in forward flight, with the lift it provides increasing with increasing air speed of the hybrid helicopter.
In order to increase the available thrust, a hybrid helicopter includes at least one propulsion arrangement, typically having at least one thrust propeller.
In addition, the thrust arrangements also provide a longitudinal stabilizing torque by being operated to deliver differential thrust. The hybrid helicopter thus has no need for an anti-torque rotor.
Furthermore, a hybrid helicopter is provided with a set of airfoils for stabilization and control purposes such as tail fins and control surfaces, in addition to the blades of the rotary wing and of the propulsion arrangements.
Another distinctive aspect of a hybrid helicopter is a proportional drive train with ratios that are constant in terms of speeds of rotation, which drive train connects together the rotary wing and the propulsion arrangements.
By modifying the pitch of the propeller blades of the propulsion arrangements collectively and by the same amount, it is possible to vary the advance thrust of the hybrid helicopter, which thrust is generated by said propulsion arrangements.
Because of the high air speeds that can be achieved by a hybrid helicopter, its airfoils for changing flight attitude are highly stressed, thus making it necessary to provide power assistance in order to enable them to be moved manually.
For this purpose, the aircraft is fitted with manual controls. The term “manual control” should not be interpreted strictly, but rather in contrast to automatic control that is actuated independently of the human action of the pilot. That is why, a rudder bar for controlling a change of attitude in yaw should be regarded as being a manual control in spite of the fact that it is actuated by a pedal unit.
Conventionally, power assistance to manual controls deprives pilots of tactile feedback (known as “haptic” feedback) that would otherwise enable them to sense the forces being applied to the airfoils that are to be moved, and thus making it possible to assess the consequences of a commanded action, and in particular whether such an action might endanger the aircraft.
Such a risk arises when the structural and/or functional limits concerning flight are reached while performing an operation for changing a flight attitude.
Thus, with power-assisted manual flight controls, it is sometimes necessary to be able to vary the gain of a manual control device such as a stick, a lever, or a rudder bar, for example.
Thus, for a given stroke applied as input to such a manual control device it is possible to obtain an output stroke that varies in proportion to some setpoint that is itself representative of a piloting parameter of the aircraft.
With a hybrid helicopter, air speed may reach values greater than 400 km/h, for example. A hybrid helicopter is thus subjected to an increased risk of exceeding its structural and/or flying limits, in the event of a manual flight control being operated too hard in the absence of any limit on its gain. One of the risks involved involves automatically taking a nose-up attitude.
Consequently, the narrow technical field of the invention is that of means for adjusting the gain of certain manual flight-attitude-changing controls of a helicopter in proportion to a piloting parameter of said helicopter.
The term “manual” flight control is used herein to cover various control devices under the authority of at least one pilot and suitable for modifying the flying attitude of a helicopter. This may be a control for acting on the cyclic pitch of the blades of a rotary wing, in pitching or in roll, generally applied using a so-called “cyclic stick”. Such manual flight controls also include controls that act on the collective pitch (angle of attack) of the blades of a rotary wing, in general by using a lever known as a “collective pitch lever” or as a collective lever.
In a modern helicopter, such manual controls are located in the cockpit of the aircraft.
In this narrow technical field of means for obtaining proportional adaptation of the gain of manual flight controls, document FR 1 132 452 describes a flight control transmission mechanism for an airplane or an aerodyne, the mechanism having a ratio that is variable in proportion as a function of air speed so as to limit automatically the movements of control surfaces as the air speed increases. The mechanism is interposed between the control on which the pilot acts and the control surface that is to be moved. A servo-mechanism such as a hydraulic servo-control serves to actuate the control surface via a linkage including an adjustable pivot axis. The resistance to movement of the control is also increased proportionally as a function of Mach number.
Document FR 2 476 013 describes a flight control device for an airplane that presents characteristics that are different in fast flight and in slow flight. That device is interposed between a pilot control and a control member to apply transmission that is continuously adjustable as a function of a variable flight factor such as a dynamic pressure. That device includes a rocking lever having two arms hinged to links, or cables, and an abutment-adjusting member in the form of a projection that is moved further away or closer as a function of the value of the variable factor.
Document FR 2 916 420 describes a hybrid type helicopter. In addition to its rotary wing, that hybrid helicopter includes a fixed wing that provides a fraction of its lift in forward flight, which fraction increases with increasing air speed of the hybrid helicopter.
In order to increase the available thrust, a hybrid helicopter includes at least one propulsion arrangement, typically having at least one thrust propeller. In addition, the thrust arrangements also provide a longitudinal stabilizing torque by being operated to deliver differential thrust. Furthermore, a hybrid helicopter is provided with a set of airfoils for stabilization and control purposes such as tail fins and control surfaces, in addition to the blades of the rotary wing and of the propulsion arrangements. Because of the high air speeds that can be achieved by a hybrid helicopter, its airfoils for changing flight attitude are highly stressed, thus making it necessary to provide power assistance in order to enable them to be moved manually. For this purpose, the aircraft is fitted with manual controls.
Document GB 613 715 describes a rotary wing aircraft having a flight control system that is adjusted as a function of variations in flight of the power applied to thrusters.
Document GB 1 180 311 describes a lever mechanism for controlling two control surfaces in opposite directions, either on an aircraft or on a watercraft.
Document U.S. Pat. No. 3,168,265 describes a variable-ratio flight control device for an aircraft capable both of forward flight and of hovering, using levers that are connected via a hinge of angle that is adjustable as a function of the aircraft's true airspeed.
Document U.S. Pat. No. 3,218,874 describes a variable-ratio flight control system for an aircraft. The ratio varies from zero to a determined value within a mechanical flight control linkage, depending on a relative angular adjustment of two levers.
Mention may also be made of document EP 1 918 196 that describes a control stick system for an airliner with force-weighted response known as “haptic” feedback. The more the speed of the airplane increases, the greater the reaction for any given configuration of the stick and of the airfoils of the airplane. The airfoils are flaps hinged to the wings of the fixed wing of the airplane. Various factors having an influence on this reaction are taken into consideration, including the amount the stick is moved.
Outside the narrow technical field of flight controls, mention is also made of document JP 63 210 454 that describes a mechanical slider device that drives another slider via a link of constant length, the two respective slideways for the sliders being substantially perpendicular.
Although of interest, those prior art techniques are in practice poorly adapted to the very particular context of adjusting the gain of certain manual flight controls in a rotary wing aircraft, in particular a hybrid helicopter, in proportion to a piloting parameter.
Document U.S. Pat. No. 3,799,695 describes control devices for a helicopter. In control linkages, certain cranks are in the form of levers, each provided with a respective slot in which a finger can slide in order to vary the lever arm of said lever as a function of detections performed by an electronic device coupled to said crank, as a function of the flying speed of the helicopter.
Document U.S. Pat. No. 3,589,331 describes a control system for a rotary wing vehicle. Mechanisms work in parallel and in unison at low flying speeds and act differently at high speeds.
With those documents mentioned, it is easier to look in depth at the technical problems on which the invention is based, relying in particular on the example of the invention being applied to a modern rotary wing aircraft, e.g. a hybrid helicopter.
At forward speeds greater than those of a conventional rotary wing aircraft, it is particularly complex in practice to provide the functional balancing needed for optimizing the flight of such a hybrid helicopter.
Thus, it is common to find that downstream from an independent linkage for changing a given flight attitude directly in response to a corresponding manual control device, imposed flight or structural parameters of the aircraft give rise to changes of attitude other than the change requested via said linkage.
For example, an independent linkage for changing flight attitude in yaw, as actuated by the pilot using a control device such as a rudder bar, gives rise to changes of attitude that are combined with those actually being initiated by the pilot, e.g. in roll.
As mentioned above, in the absence of power assistance, the reaction forces on the airfoils give rise to tactile sensations (haptic feedback) representative of certain stresses to which the aircraft is subjected.
With power assistance of manual controls, the absence of tactile sensation representative of these reaction forces means that the pilot cannot perceive a risk of danger when the functional or structural limits of the aircraft are close or even exceeded.
Furthermore, it is found that the forces imposed on the airfoils of a modern helicopter are a function of its instantaneous flight conditions, and in particular of its load factor and its air speed.
The maximum load factor that can be accepted by an aircraft while turning or pulling up is proportional to said instantaneous air speed and also to its instantaneous angular velocity about the turning or pull-up axis. For example, the angular velocity in yaw of an aircraft is the rate at which it turns in one direction or the other about an axis extending in the elevation direction, when turning to left or to right.
In order to avoid exceeding a limiting load factor, a manual flight control is often provided with its own force relationship so as to indicate more clearly when such a limiting load factor is close.
If a manual flight control, e.g. for controlling yaw, is actuated at low frequencies, e.g. by slowly moving the corresponding control device, then the aircraft responds in a mode that is practically steady, i.e. without any variation in the corresponding attitude. As a result, the functional or structural limits of the aircraft are generally not reached. Conversely, if the same manual flight control is actuated at high frequency, i.e. by moving the control device quickly, then the fuselage, so to speak, does not have enough time to keep up.
Under such conditions, the functional or structural limits of the aircraft run the risk of being reached. Departures in one or more of said flight attitudes have a tendency to increase in significant manner and can lead to situations that are unacceptable, or indeed dangerous such as a tendency to take a nose-up attitude.
Clearly, those phenomena are more penalizing or even dangerous with increasing air speed of the aircraft.
In an attempt to avoid such situations in flight, the movements of certain control devices, in particular in yaw, need to have gain that is limited appropriately, in a manner that is proportional to a flight parameter that is representative of the corresponding risk.
In other words, it is desirable to vary the sensitivity of the flight-attitude-changing linkage controlled by a manual control, thereby reducing the pilot's authority over the airfoil situated at the end of said linkage on approaching a limiting value for a load factor.
Nevertheless, it is appropriate to ensure that such a variation in the sensitivity of the linkage does not give rise in interfering manner and by reversibility to modifications in the gains of certain constituent parts of said linkage.
Although various techniques have conventionally been used for varying the sensitivity of manual controls in this way, they are found not to be entirely satisfactory in practice, in particular for a fast rotary wing aircraft.
In order to reach the invention, particular testing has been performed. That testing has shown that within a given modern aircraft, optimizing the gain of the manual flight control device can be performed as a function of the instantaneous position of a manual flight control member that is distinct from the flight control device of gain that is to be adjusted. In particular, such a control member is one that acts on the air speed of the aircraft.
Furthermore, in order to increase the reliability of an aircraft, it has been found advantageous to minimize redundancies between an automatic electronic flight installation suitable for performing a given flight attitude change, e.g. an autopilot, and a control linkage acting on the same change of attitude.
Thus, in certain circumstances, it has been found undesirable for the optimization or adjustment of the gain of a given manual flight control device to depend directly on the electronic flight installation. Attempts have therefore been made to achieve said optimization within said linkage dedicated to the change of attitude in question in purely mechanical manner.
Nevertheless, those approaches do not rule out logical, electrical, and/or hydraulic processing for optimizing gain, damping, haptic feedback, and the like, acting on such and such a manual control device in a given helicopter.
In this context, it has been found that a bellcrank of the kind that is conventionally to be found in a flight-attitude-changing linkage, is a component that is suitable in certain applications for incorporating mechanical adjustment of the gain of a manual control device such as a rudder bar, a lever, a knob, or a stick.
It is common practice for one or more bellcranks to exist in a given flight-attitude-changing linkage of a modern aircraft. Conventionally, a bellcrank is operated by a manual control device such as a rudder bar when providing yaw control.